This invention relates generally to turbine engine nozzles and more particularly, to methods and apparatus for securing multi-piece nozzle assemblies.
At least some known turbine engines include a turbine nozzle assembly which channels flow towards a turbine. At least some known turbine nozzle assemblies include a plurality of nozzles arranged circumferentially within the engine. Each nozzle includes an airfoil vane that extends between inner and outer band platforms. Each airfoil vane includes a pair of sidewalls that are connected at a leading edge and a trailing edge.
During operation, the nozzles are typically cooled by a combination of internal convective cooling and gas side film cooling. Typically, the metal temperature distribution of a vane airfoil is such that the trailing edge is significantly hotter than a temperature of the bulk of the airfoil. The temperature gradient created may induce compressive stresses at the vane trailing edge. The combination of such stresses and temperatures may result in the vane trailing edge being the life limiting location of the nozzle.
The overall efficiency of the gas turbine engine is directly related to the temperature of the combustion gases, and as such, engine efficiency may be limited by the ability to operate the turbine nozzle at high temperature. As such, cooling engine components, including the turbine components, is necessary to facilitate reducing thermal stresses induced to such components. Accordingly, at least some known turbine nozzles include cavity cooling circuits which define flow paths for channeling cooling air flow through the cavity for cooling the airfoil, prior to the air flow being discharged downstream through trailing edge slots defined within the airfoil. Because of material limitations, known nozzle airfoils may require a complex cooling scheme to reduce operating temperatures within the airfoil.